Gas turbine engine and method of operating

ABSTRACT

A gas turbine engine includes; a compressor, a combustor, and a turbine in serial flow relationship; a heat exchanger, the heat exchanger having an inlet, an outlet, and an internal surface coated with a catalyst, the heat exchanger being located upstream of the compressor; a source of hydrocarbon fuel in fluid communication with the inlet of the heat exchanger; a source of oxygen in fluid communication with the inlet of the heat exchanger; and a distribution system for receiving reformed hydrocarbon fuel from the heat exchanger.

PRIORITY INFORMATION

The present disclosure claims priority to U.S. Patent Application Ser.No. 62/990,174 filed on Mar. 16, 2020, which is incorporated byreference herein in its entirety.

FIELD OF TECHNOLOGY

The present invention pertains to aircraft components, such as gasturbine engines, and more particularly to those which are exposed tohigh levels of heat in operation, such as during high-speed flight.

BACKGROUND OF THE INVENTION

Many aircraft components are exposed to heat sources during operation,such as internal components of propulsion systems including gas turbineengines. Particularly during high-speed flight operations, thesecomponents as well as external components of the aircraft itself areexposed to heat from skin friction due to the speed with which theaircraft is traveling through the atmosphere. Components such as theaircraft skin, leading edges of structures such as wings, chines, andcontrol surfaces, and engine inlets may be particularly affected inaddition to the temperature of the air entering a gas turbine engine.

These heat sources may cause surface and internal temperatures of suchaircraft components to exceed their structural and/or operationalcapabilities, requiring expensive or exotic materials and changes inoperating characteristics.

It would therefore be desirable to provide a system and method forremoving heat from aircraft components which is reliable and durable inoperational service and capable of removing heat from such componentsduring high-speed flight.

It would be further desirable to capture the heat for use in improvingthe propulsive efficiency of the aircraft.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, a gas turbine engine includes; a compressor, a combustor,and a turbine in serial flow relationship; a heat exchanger, the heatexchanger having an inlet, an outlet, and an internal surface coatedwith a catalyst, the heat exchanger being located upstream of thecompressor; a source of hydrocarbon fuel in fluid communication with theinlet of the heat exchanger; a source of oxygen in fluid communicationwith the inlet of the heat exchanger; and a distribution system forreceiving reformed hydrocarbon fuel from the heat exchanger.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedFigures, in which:

FIG. 1 is a perspective illustration of an exemplary embodiment of ahigh-speed aircraft suitable for implementing the heat removal apparatusand methods described herein;

FIG. 2 is a cross-sectional schematic illustration of a tube of anexemplary heat exchanger as described herein;

FIG. 3 is a cross-sectional schematic illustration of an exemplaryembodiment of a heat exchanger as described herein having an array of aplurality of tubular structures;

FIG. 4 is a plan view schematic illustration of an exemplary embodimentof a heat exchanger as described herein, having an array of a pluralityof tubular structures and depicting inlet and outlet manifolds;

FIG. 5 is a cross-sectional schematic illustration of an exemplaryembodiment of a gas turbine engine suitable for use as an aircraftpropulsion system and incorporating a heat exchanger as describedherein;

FIG. 6 is a schematic illustration of a gas turbine engine utilizing aheat exchanger as described herein downstream of the last turbine stageto reform a hydrocarbon fuel;

FIG. 7 is a schematic illustration of a gas turbine engine, similar toFIG. 6, utilizing a heat exchanger as described herein between turbinestages to reform a hydrocarbon fuel;

FIG. 8 is a schematic diagram illustrating a method of operating theexemplary embodiment of FIG. 5;

FIG. 9 is a cross-sectional schematic illustration of an exemplaryembodiment of a gas turbine engine suitable for use as an aircraftpropulsion system similar to the embodiment of FIG. 5, but adapted toinclude a hyperburner system; and

FIG. 10 is a schematic illustration of an exemplary embodiment of a heatexchanger system which incorporates a plurality of heat exchangerssimilar to the heat exchanger of FIG. 4, the system being configured tooperate with the gas turbine engine of FIG. 9.

Corresponding reference characters indicate corresponding partsthroughout the several views. The exemplifications set out hereinillustrate exemplary embodiments of the disclosure, and suchexemplifications are not to be construed as limiting the scope of thedisclosure in any manner.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention.

The following description is provided to enable those skilled in the artto make and use the described embodiments contemplated for carrying outthe invention. Various modifications, equivalents, variations, andalternatives, however, will remain readily apparent to those skilled inthe art. Any and all such modifications, variations, equivalents, andalternatives are intended to fall within the spirit and scope of thepresent invention.

All directional references (e.g., radial, axial, proximal, distal,upper, lower, upward, downward, left, right, lateral, front, back, top,bottom, above, below, vertical, horizontal, clockwise, counterclockwise,upstream, downstream, forward, aft, etc.) are only used foridentification purposes to aid the reader's understanding of the presentinvention, and do not create limitations, particularly as to theposition, orientation, or use of the invention. Connection references(e.g., attached, coupled, connected, and joined) are to be construedbroadly and can include intermediate members between a collection ofelements and relative movement between elements unless otherwiseindicated. As such, connection references do not necessarily infer thattwo elements are directly connected and in fixed relation to oneanother. The exemplary drawings are for purposes of illustration onlyand the dimensions, positions, order and relative sizes reflected in thedrawings attached hereto can vary.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise. Moreover, the suffix “(s)” asused herein is usually intended to include both the singular and theplural of the term that it modifies, thereby including one or more ofthat term.

As used herein, the term “or” is not meant to be exclusive and refers toat least one of the referenced components (for example, a material)being present and includes instances in which a combination of thereferenced components may be present, unless the context clearlydictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 10percent margin.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

As used herein, the terms “may” and “may be” indicate a possibility ofan occurrence within a set of circumstances; a possession of a specifiedproperty, characteristic or function; and/or qualify another verb byexpressing one or more of an ability, capability, or possibilityassociated with the qualified verb. Accordingly, usage of “may” and “maybe” indicates that a modified term is apparently appropriate, capable,or suitable for an indicated capacity, function, or usage, while takinginto account that in some circumstances, the modified term may sometimesnot be appropriate, capable, or suitable. For example, in somecircumstances, an event or capacity can be expected, while in othercircumstances, the event or capacity cannot occur. This distinction iscaptured by the terms “may” and “may be”.

Reference throughout the specification to “some embodiments”, and soforth, means that a particular element (e.g., feature, structure, and/orcharacteristic) described in connection with the invention is includedin at least one embodiment described herein, and may or may not bepresent in other embodiments. In addition, it is to be understood thatthe described inventive features may be combined in any suitable mannerin the various embodiments.

Various aspects of the invention are explained more fully with referenceto the exemplary embodiments discussed below. It should be understoodthat, in general, the features of one embodiment also may be used incombination with features of another embodiment, and that theembodiments are not intended to limit the scope of the invention.

FIG. 1 perspective illustration of an exemplary embodiment of ahigh-speed aircraft 10 suitable for implementing the heat removalapparatus and methods described herein. As used herein, the term“high-speed aircraft” is intended to refer to aircraft designed tooperate above the speed of sound, i.e., above Mach 1, and moreparticularly to aircraft designed to operate in hypersonic flightregimes above Mach 5, such as in the range of Mach 5 to Mach 10.

In the configuration shown in FIG. 1, the exemplary high-speed aircraft10 includes a fuselage 11, wings 12, vertical stabilizers 13, leadingedges 14 of the wings 12 and fuselage 11, as well as gas turbine engines15 serving as aircraft propulsion systems. An outer skin surface 16covers at least portions of the fuselage 11 and wings 12. In high-speedflight, movement of the aircraft 10 through the atmosphere causesheating of aircraft surfaces such as the skin surface 16 on the exteriorof the aircraft 10, and in particular causes heat buildup in the regionsof the leading edges 14 of the wings 12 and fuselage 11. High-speedoperation also causes the gas turbine engines 15 to experience hightemperature operation as atmospheric air is reduced to subsonic speedswhich increases pressures within the gas turbine engine. Each of theselocations may utilize a system for removing heat from aircraftcomponents as described herein. Systems of interest are furtherdescribed in commonly-assigned, co-pending U.S. patent application Ser.No. 16/681,292, filed Nov. 12, 2019, the specification and drawings ofwhich are included as Appendix 1 and Appendix 2, respectively.

FIG. 2 is a cross-sectional schematic illustration of a tube of anexemplary heat removal system 20 in the form of a heat exchanger asdescribed herein. Heat removal system 20 includes at least one tubularstructure 21 having an inner surface 22 and an outer surface 23. Theinner surface 22 includes a coating 24, as will be described hereafter,and an interior space 25 located inwardly of the interior surface 22 andthe coating 24 through which a fluid may pass.

The inner surface of the tubular structure 21, which is accessible to afluid such as a hydrocarbon fuel in the interior space 25, comprises acoating 24 of perovskite material and the tuning material.

As used herein the term “hydrocarbon cracking”, “cracking hydrocarbon”,or any variation thereof, refers to but is not limited to processes inwhich hydrocarbons are cracked in apparatuses to obtain materials withsmaller molecules. The hydrocarbon may include ethane, heptane, liquidpetroleum gas, naphtha, gas oil, bottoms from atmospheric and vacuumdistillation of crude oil, or any combination thereof.

As used herein the term “coke” or any variation thereof refers to but isnot limited to carbonaceous solid or liquid, or particulates ormacromolecules forming the carbonaceous solid or liquid, which arederived from coal, petroleum, wood, hydrocarbons and other materialscontaining carbon.

As used herein the term “perovskite material” or any variation thereofrefers to but is not limited to any material having an ABO3 perovskitestructure and being of formula A_(a)B_(b)O_(3−δ), wherein 0.9<a≤1.2;0.9<b≤1.2; −0.5<δ<0.5; A comprises a first element and optionally asecond element, the first element is selected from calcium (Ca),strontium (Sr), barium (Ba), lithium (Li), sodium (Na), potassium (K),rubidium (Rb) and any combination thereof, the second element isselected from yttrium (Y), bismuth (Bi), lanthanum (La), cerium (Ce),praseodymium (Pr), neodymium (Nd), promethium (Pm), samarium (Sm),europium (Eu), gadolinium (Gd), terbium (Tb), dysprosium (Dy), holmium(Ho), erbium (Er), thulium (Tm), ytterbium (Yb), lutetium (Lu) and anycombination thereof; and B is selected from silver (Ag), gold (Au),cadmium (Cd), cerium (Ce), cobalt (Co), chromium (Cr), copper (Cu),dysprosium (Dy), erbium (Er), europium (Eu), ferrum (Fe), gallium (Ga),gadolinium (Gd), hafnium (Hf), holmium (Ho), indium (In), iridium (Ir),lanthanum (La), lutetium (Lu), manganese (Mn), molybdenum (Mo), niobium(Nb), neodymium (Nd), nickel (Ni), osmium (Os), palladium (Pd),promethium (Pm), praseodymium (Pr), platinum (Pt), rhenium (Re), rhodium(Rh), ruthenium (Ru), antimony (Sb), scandium (Sc), samarium (Sm), tin(Sn), tantalum (Ta), terbium (Tb), technetium (Tc), titanium (Ti),thulium (Tm), vanadium (V), tungsten (W), yttrium (Y), ytterbium (Yb),zinc (Zn), zirconium (Zr), and any combination thereof.

In some embodiments, the perovskite material may be of formulan(A_(a)B_(b)O_(3−δ)), in which n=2, 3, 4, 8, and etc., and the formulaA_(a)B_(b)O_(3−δ) is the simplified form thereof. In some embodiments,in the ABO₃ perovskite structure, A cations are surrounded by twelveanions in cubo-octahedral coordination, B cations are surrounded by sixanions in octahedral coordination and oxygen anions are coordinated bytwo B cations and four A cations. In some embodiments, the ABO₃perovskite structure is built from corner-sharing BO₆ octahedra. In someembodiments, the ABO₃ perovskite structure includes distortedderivatives. The distortions may be due to rotation or tilting ofregular, rigid octahedra or due to the presence of distorted BO₆octahedra. In some embodiments, the ABO₃ perovskite structure is cubic.In some embodiments, the ABO₃ perovskite structure is hexagonal.

In some embodiments, A only comprises the first element. The firstelement may be a single element or a combination of elements selectedfrom calcium (Ca), strontium (Sr), barium (Ba), lithium (Li), sodium(Na), potassium (K), and rubidium (Rb).

In some embodiments, A comprises a combination of the first element andthe second element. The second element may be a single element or acombination of elements selected from yttrium (Y), bismuth (Bi),lanthanum (La), cerium (Ce), praseodymium (Pr), neodymium (Nd),promethium (Pm), samarium (Sm), europium (Eu), gadolinium (Gd), terbium(Tb), dysprosium (Dy), holmium (Ho), erbium (Er), thulium (Tm),ytterbium (Yb), and lutetium (Lu).

Likewise, B may be a single element or a combination of elementsselected from silver (Ag), gold (Au), cadmium (Cd), cerium (Ce), cobalt(Co), chromium (Cr), copper (Cu), dysprosium (Dy), erbium (Er), europium(Eu), ferrum (Fe), gallium (Ga), gadolinium (Gd), hafnium (Hf), holmium(Ho), indium (In), iridium (Ir), lanthanum (La), lutetium (Lu),manganese (Mn), molybdenum (Mo), niobium (Nb), neodymium (Nd), nickel(Ni), osmium (Os), palladium (Pd), promethium (Pm), praseodymium (Pr),platinum (Pt), rhenium (Re), rhodium (Rh), ruthenium (Ru), antimony(Sb), scandium (Sc), samarium (Sm), tin (Sn), tantalum (Ta), terbium(Tb), technetium (Tc), titanium (Ti), thulium (Tm), vanadium (V),tungsten (W), yttrium (Y), ytterbium (Yb), zinc (Zn), and zirconium(Zr).

In some embodiments, the perovskite material comprises SrCeO₃,SrZr_(0.3)Ce_(0.7)O₃, BaMnO₃, BaCeO₃, BaZr_(0.3)Ce_(0.7)O₃,BaZr_(0.3)Ce_(0.5)Y_(0.2)O₃, BaZr_(0.1)Ce_(0.7)Y_(0.2)O₃, BaZrO₃,BaZr_(0.7)Ce_(0.3)O₃, BaCe_(0.5)Zr_(0.5)O₃, BaCe_(0.9)Y_(0.1)O₃,BaCe_(0.85)Y_(0.15)O₃, or BaCe_(0.8)Y_(0.2)O₃. For example, for SrCeO₃,A is Sr, a=1, B is Ce, b=1, and δ=0. For SrZr_(0.3)Ce_(0.7)O₃, A is Sr,a=1, B is a combination of Zr and Ce, b=1, and 6=0. For BaMnO₃, A is Ba,a=1, B is Mn, b=1, and 6=0. For BaCeO₃, A is Ba, a=1, B is Ce, b=1, and6=0. For BaZr_(0.3)Ce_(0.7)O₃, A is Ba, a=1, B is a combination of Zrand Ce, b=1, and δ=0. For BaZr_(0.3)Ce_(0.5)Y_(0.2)O₃, A is Ba, a=1, Bis a combination of Zr, Ce and Y, b=1, and δ=0.

In some embodiments, the perovskite material comprisesLa_(0.1)Ba_(0.9)Ce_(0.7)Zr_(0.2)Y_(0.1)O₃,Ce_(0.1)Ba_(0.9)Ce_(0.7)Zr_(0.2)Y_(0.1)O_(3.05),Ce_(0.5)Ba_(0.5)Ce_(0.7)Zr_(0.2)Y_(0.1)O_(3.45),Y_(0.1)Ba_(0.9)Ce_(0.7)Zr_(0.2)Y_(0.1)O₃,Y_(0.5)Ba_(0.5)Ce_(0.7)Zr_(0.2)Y_(0.1)O_(3.2),Bi_(0.1)Ba_(0.9)Ce_(0.7)Zr_(0.2)Y_(0.1)O₃,Bi_(0.5)Ba_(0.5)Ce_(0.7)Zr_(0.2)Y_(0.1)O_(3.2),Pr_(0.1)Ba_(0.9)Ce_(0.7)Zr_(0.2)Y_(0.1)O₃, orPr_(0.5)Ba_(0.5)Ce_(0.7)Zr_(0.2)Y_(0.1)O_(3.2). ForLa_(0.1)Ba_(0.9)Ce_(0.7)Zr_(0.2)Y_(0.1)O₃, A is a combination of Ba andLa, the first element is La, the second element is Ba, a=1, B is acombination of Ce, Zr and Y, b=1, and, δ=0. ForCe_(0.1)Ba_(0.9)Ce_(0.7)Zr_(0.2)Y_(0.1)O_(3.05) andCe_(0.5)Ba_(0.5)Ce_(0.7)Zr_(0.2)Y_(0.1)O_(3.45), A is a combination ofCe and Ba, the first element is Ce, the second element is Ba, a=1, B isa combination of Ce, Zr and Y, b=1, and, δ=−0.05 and −0.45,respectively. For Y_(0.1)Ba_(0.9)Ce_(0.7)Zr_(0.2)Y_(0.1)O₃ andY_(0.5)Ba_(0.5)Ce_(0.7)Zr_(0.2)Y_(0.1)O_(3.2), A is a combination of Yand Ba, the first element is Y, the second element is Ba, a=1, B is acombination of Ce, Zr and Y, b=1, and, δ=0 and −0.2, respectively. ForBi_(0.1)Ba_(0.9)Ce_(0.7)Zr_(0.2)Y_(0.1)O₃ andBi_(0.5)Ba_(0.5)Ce_(0.7)Zr_(0.2)Y_(0.1)O_(3.2), A is a combination of Biand Ba, the first element is Bi, the second element is Ba, a=1, B is acombination of Ce, Zr and Y, b=1, and, δ=0 and −0.2, respectively.Similarly, for Pr_(0.1)Ba_(0.9)Ce_(0.7)Zr_(0.2)Y_(0.1)O₃ andPr_(0.5)Ba_(0.5)Ce_(0.7)Zr_(0.2)Y_(0.1)O_(3.2), A is a combination of Prand Ba, the first element is Pr, the second element is Ba, a=1, B is acombination of Ce, Zr and Y, b=1, and, δ=0 and −0.2, respectively.

In some embodiments, the perovskite material comprisesBaZr_(0.3)Ce_(0.7)O₃.

As used herein the term “tuning material” or any variation thereofrefers to any material that reduces the yield of carbon monoxide inhydrocarbon cracking. The tuning material may comprise one material or acombination of multiple materials. In some embodiments, the tuningmaterial comprises zirconium oxide, doped zirconium oxide, or anyprecursor or combination thereof.

In some embodiments, the method for cracking hydrocarbon is operated ata temperature in a range from about 700° C. to about 900° C. with thepresence of steam, a weight ratio of steam to hydrocarbon is in a rangefrom about 3:7 to about 7:3, and the hydrocarbon includes ethane,heptane, liquid petroleum gas, naphtha, gas oil, or any combinationthereof.

In some embodiments, the method for cracking hydrocarbon is operated ata temperature in a range from about 480° C. to about 600° C. in thepresence of steam, the hydrocarbon comprises bottoms from atmosphericand vacuum distillation of crude oil and a weight percentage of steam isin a range from about 1 wt % to about 2 wt %.

The perovskite material may or may not chemically react with the tuningmaterial. Thus, the inner surface may comprise a combination or areaction product of the perovskite material and the tuning material. Insome embodiments, the inner surface comprises a combination of theperovskite material, the tuning material and a reaction product of theperovskite material and the tuning material.

The perovskite material and the tuning material may be in a coatingapplied to the apparatus using different methods, for example, airplasma spray, slurry coating, so-gel coating, and solution coating. Insome embodiments, the perovskite material and the tuning material arecoated using slurry coating method.

The amount of the tuning material and the perovskite material in theslurry may vary as long as a continuous, strong, carbon monoxidereducing and anticoking coating is formed, depending on the specifictuning material and the perovskite material being used and the workingcondition of the coating. In some embodiments, a weight ratio of theperovskite material to the tuning material is from about 7:3 to about7:93. In some embodiments, a weight of the perovskite material is equalto or less than that of the tuning material.

The slurry may further comprise an organic binder, an inorganic binder,a wetting agent, a solvent or any combination thereof to enhance theslurry wetting ability, tune the slurry viscosity or get good greencoating strength. When the organic binder, the inorganic binder, thewetting agent, the solvent or any combination thereof is added in theslurry, a total weight percentage of the tuning material and theperovskite material in the slurry may be from about 10% to about 90%, orpreferably from about 15% to about 70%, or more preferably from about30% to about 55%.

In some embodiments, the slurry comprises the perovskite material, thetuning material, cerium oxide, yttrium oxide, glycerol, and polyvinylalcohol (PVA).

The slurry may be applied to the apparatus by different techniques, suchas at least one of sponging, painting, centrifuging, spraying, fillingand draining, and dipping. In some embodiments, the slurry is applied bydipping, i.e., dipping the part to be coated in the slurry. In someembodiments, the slurry is applied by filling and draining, i.e.,filling the slurry in the article to be coated and draining out theslurry afterwards by, e.g., gravity.

Additional descriptions of the cracking methods and systems may be foundin published patent documents, all of which are incorporated byreference: U.S. Pat. No. 9,499,747, WO2015105589A1, CA2821249A1,US20170260460, CA2932461A1, WO2015088671A1, and US20170022428.

After the slurry is applied to the apparatus, a sintering process may befollowed. As used herein the term “sintering” or any variation thereofrefers to but is not limited to a method of heating the material in asintering furnace or other heater facility. In some embodiments, thesintering temperature is in a range from about 850° C. to about 1700° C.In some embodiments, the sintering is at about 1000° C. In someembodiments, the sintering is carried out in an inert atmosphere such asargon or nitrogen. In some embodiments, the sintering is preceded by aheat treatment in air to form an oxide layer on the inner surface of thetube that improves coating adhesion.

In operation, a hydrocarbon such as an aircraft fuel is fed into theinterior space 25 along with an oxygen-containing species, such assteam, which may be provided in the form of liquid water which vaporizesin the presence of sufficient heat, or oxygenated fuels such as ethanolor methanol. The cracking of the hydrocarbon fuel which takes placewithin the interior space 25 is an endothermic reaction in which all thecarbon-carbon bonds are broken and hydrogen and methylene radicals areformed. This highly endothermic process takes heat away from, andtherefore cools, the tubular structure 21 and its surroundingenvironment. The cracking of the hydrocarbon therefore transformstubular structure 21 into a heat exchanger 20 and functions as a heatremoval system which can be employed to remove heat from components 30of a high-speed aircraft 10. The coating reduces or prevents theformation of coke in the interior space 25, which could eventuallyimpede the flow of fuel and reduce the capability of the heat exchanger20. The reformed fuel exiting the heat exchanger 20 may then be utilizedas fuel for aircraft propulsion.

FIG. 3 is a cross-sectional schematic illustration of an exemplaryembodiment of a heat removal system in the form of heat exchanger 20 asdescribed herein, having an array of a plurality of tubular structures21, each with an interior space 25. The heat exchanger 20 is located inclose proximity to an aircraft component 30, from which it is desired toremove thermal energy (heat). Heat exchanger 20 may be joined to, orintegrally formed with, the aircraft component through conventional oradditive manufacturing techniques known in the art. Alternatively, heatexchanger 20 may be employed as an air-to-fuel heat exchanger and thusutilized to remove heat from an air stream flowing through the array oftubular structures 21.

FIG. 4 is a plan view schematic illustration of an exemplary embodimentof a heat exchanger 20 as described herein, having an array of aplurality of tubular structures 21 and depicting inlet and outletmanifolds, 26 and 28, respectively. Manifolds 26 and 28, respectively,fluidly couple inlet and outlet ends of the tubular structures 21 tocommon inlet 27 and outlet 29. Inlet 27 may be in turn fluidly coupledto a pipe or conduit which is a source of hydrocarbon fuel and/or steam(water). Outlet 29 may be in turn fluidly coupled to a pipe or conduitwhich is a destination or recipient of reformed fuel after crackingtakes place within the tubular structures 21 of the heat exchanger 20.

A heat exchanger as described herein may in fact comprise a plurality ofheat exchangers separately manifolded, in addition to a plurality oftubular structures sharing a common manifold. A control system,comprising sensors, valves, and/or electronic control actuators maycontrol flow through individual tubular structures and/or between andamong a plurality of heat exchangers. This may provide flexibility fordifferent operating conditions as well as to cycle between heatexchangers if it is necessary or desirable to take some tubularstructures or some heat exchangers offline to remove any coke depositswhich may accumulate during operation.

FIG. 5 is a cross-sectional schematic illustration of an exemplaryembodiment of a gas turbine engine 15 suitable for use as an aircraftpropulsion system for high-speed aircraft 10 and incorporating a heatexchanger 20 as described herein. As shown in FIG. 5, the heat exchanger20, containing catalyst coating 24 within tubular structures 21, islocated within the inlet section 31 of the gas turbine engine 15. Thecasing surrounding the inlet section serves as the aircraft component 30which is it desired to remove heat from which is generated by andtransported by the incoming air stream 33 during high-speed flight. Thislowers the temperature of the air prior to entering the compressorsection 32 of the gas turbine engine 15. The reformed fuel exiting theheat exchanger 20 may then be fed into the combustor section 34 of thegas turbine engine 15.

FIG. 6 is a schematic illustration of a gas turbine engine 15 utilizinga heat exchanger 20 as described herein downstream of the last turbinestage, aft of the high pressure turbine 35 and low pressure turbine 36,to reform a hydrocarbon fuel. The heat contained in the residual streamof air exiting the low pressure turbine 36 provides the energy to crackthe fuel and is then exhausted to the atmosphere 37. As discussedpreviously, the reformed fuel may be fed into the combustor 34 to fuelthe gas turbine engine 15.

FIG. 7 is a schematic illustration of a gas turbine engine 15, similarto FIG. 6, utilizing a heat exchanger 20 as described herein between thehigh and low pressure turbine stages 35 and 36 to reform a hydrocarbonfuel.

With any of the exemplary embodiments described herein, a plurality ofheat removal systems may be employed either in series or in parallel,and may share inlets and outlets or may be separately plumbed withindividual sources and destinations for hydrocarbon fuel and reformedfuel.

With parallel heat removal systems, all systems may be operatedconcurrently or some systems may be deactivated for rejuvenation andremoval of coke deposits or to modulate the level of heat removalcapacity during various phases of aircraft operation.

FIG. 8 is a schematic diagram illustrating a method of operating theexemplary embodiment of FIG. 5. In FIG. 8, hot inlet air 33 enters thefront of the gas turbine engine 15. The heat exchanger 20, which may bea single element arranged annularly and/or symmetrically about theengine centerline 40 or a plurality of individual heat exchangerelements, receives a liquid fuel 41 through inlet 27. This liquid fuel41 then becomes vaporized fuel 42 which is then combined withoxygen-containing medium, such as bleed air 44 from the gas turbineengine compressor or water from a storage tank, for example, within theheat exchanger 20. The heat exchanger 20 then reforms the fuel using theavailable heat from the hot inlet air 33 and the oxygen-containingmedium, such that reformed fuel 43 then exits the heat exchanger 20through outlet 29. Conditioned air 45, which has a lower temperaturethan the hot inlet air 33, then flows to the gas turbine enginecompressor 32, in the embodiments of FIGS. 5 and 9, or to the atmospherein the embodiment of FIG. 6, or to the low pressure turbine 37 in theembodiment of FIG. 7. The hot inlet air 33 in the embodiment of FIGS. 5and 9 could also be exhaust from the high pressure turbine 35 in theembodiment of FIG. 7, depending upon the installed location of the heatexchanger 20.

FIG. 9 is a cross-sectional schematic illustration of an exemplaryembodiment of a gas turbine engine 15 suitable for use as an aircraftpropulsion system similar to the embodiment of FIG. 5, but adapted toinclude a hyperburner system 70. In the embodiment of FIG. 9, thehyperburner 70 is similar to an augmentor or afterburner but is designedspecifically for high-speed and/or hypersonic travel.

As with FIG. 4, in FIG. 9 the heat exchanger 20 is located in the inletsection 31 of the gas turbine engine 15. Liquid hydrocarbon fuel 41 isfed from a storage tank 52 through a supply line 58 to a first flowdivider 48, which functions as a control valve, where a portion of theliquid hydrocarbon fuel 41 may be fed to fuel nozzles 54 located withinthe combustor section 34 of the gas turbine engine 15. Liquid fuel 41may flow from the first flow divider 48 to a second flow divider, orcontrol valve, 50, where liquid fuel 41 may be directed to the heatexchanger 20 through inlet 27 via supply line 58 or selectively to fuelnozzles 56 located aft of the high and low pressure turbines 35 and 36,respectively, in the hyperburner 70. A control valve 46 controls theflow of compressor discharge air, or bleed air, 44, to the reformingcatalyst in the heat exchanger 20. Reformed (gaseous) fuel 60 departsthe heat exchanger 20 through outlet 29 and flows to the hyperburnerfuel nozzles 56 located in the hyperburner 70.

The hyperburner fuel nozzles 56 may be configured for “dual fuel”operation, that is, to operate with either liquid hydrocarbon fuel orgaseous reformed fuel, or both, as desired for the particular operatingconditions present. The control valves or flow dividers may be operatedby a control system which selectively directs liquid fuel to the gasturbine combustor and/or the hyperburner and also controls the flow ofreformed fuel to the hyperburner from the heat exchanger. Theproportions of the fuel flow to the main gas turbine engine combustorand hyperburner may be adjusted as needed, and in high-speed orhypersonic flight regimes flow splits such as about 5% to the main gasturbine engine combustor and 95% to the hyperburner may be utilized. Lowpressures, such as 3-5 bar, may be experienced in the hyperburner whichpermits the hyperburner fuel supply system to operate at low pressuresas well, minimizing the risks of leakage. Higher pressures may beexperienced in the gas turbine combustor, requiring liquid fuel to befed at higher pressures in that system.

FIG. 10 is a schematic illustration of an exemplary embodiment of a heatexchanger system 20 which incorporates a plurality of heat exchangers20A, 20B, similar to the heat exchanger 20 of FIG. 4, the system beingconfigured to operate with the gas turbine engine 15 of FIG. 9. In theembodiment of FIG. 10, the plurality of heat exchangers, or heatexchanger branches, 20A and 20B, may be separately supplied with liquidhydrocarbon fuel 41 and oxygen-containing compressor discharge air 44via control valves 71, and similar control valves 71 may direct reformedfuel to either the gas turbine engine combustor fuel nozzles 54 or thehyperburner fuel nozzles 56 located at the turbine exit. As describedpreviously, the plurality (which may be two or more) of heat exchangersmay be operated together, or one or more may be deactivated for cleaningor based on a lower demand for reformed fuel.

Various hydrocarbon fuels may be utilized with the exemplary embodimentsas described herein, including aircraft jet fuels such as Jet-A, JP-4,and JP-8, gasolines, kerosenes, rocket propulsion fuels such as RPS1 andRPS2, Diesel fuels such as D2 and D4, and blends, mixtures, andcombinations thereof.

All publications, patents and patent applications cited herein, whethersupra or infra, are hereby incorporated by reference in their entiretyto the same extent as if each individual publication, patent or patentapplication was specifically and individually indicated as incorporatedby reference. It should be appreciated that any patent, publication, orother disclosure material, in whole or in part, that is said to beincorporated by reference herein is incorporated herein only to theextent that the incorporated material does not conflict with existingdefinitions, statements, or other disclosure material set forth in thisdisclosure. As such, and to the extent necessary, the disclosure asexplicitly set forth herein supersedes any conflicting materialincorporated herein by reference. Any material, or portion thereof, thatis said to be incorporated by reference herein, but which conflicts withexisting definitions, statements, or other disclosure material set forthherein, will only be incorporated to the extent that no conflict arisesbetween that incorporated material and the existing disclosure material.

It must be noted that, as used in this specification and the appendedclaims, the singular forms “a,” “an” and “the” include plural referentsunless the content clearly dictates otherwise.

Unless defined otherwise, all technical and scientific terms used hereinhave the same meaning as commonly understood by one of ordinary skill inthe art to which the invention pertains. Although a number of methodsand materials similar or equivalent to those described herein can beused in the practice of the present invention, materials and methodsaccording to some embodiments are described herein.

It should be noted that, when employed in the present disclosure, theterms “comprises,” “comprising,” and other derivatives from the rootterm “comprise” are intended to be open-ended terms that specify thepresence of any stated features, elements, integers, steps, orcomponents, and are not intended to preclude the presence or addition ofone or more other features, elements, integers, steps, components, orgroups thereof.

As required, detailed embodiments of the present invention are disclosedherein; however, it is to be understood that the disclosed embodimentsare merely exemplary of the invention, which may be embodied in variousforms. Therefore, specific structural and functional details disclosedherein are not to be interpreted as limiting, but merely as a basis forthe claims and as a representative basis for teaching one skilled in theart to variously employ the present invention in virtually anyappropriately detailed structure.

Various characteristics, aspects, and advantages of the presentdisclosure may also be embodied in any permutation of Aspects of thedisclosure, including but not limited to the following technicalsolutions as defined in the enumerated Aspects:

1. In one aspect, a gas turbine engine includes; a compressor, acombustor, and a turbine in serial flow relationship; a heat exchanger,the heat exchanger having an inlet, an outlet, and an internal surfacecoated with a catalyst, the heat exchanger being located upstream of thecompressor; a source of hydrocarbon fuel in fluid communication with theinlet of the heat exchanger; a source of oxygen in fluid communicationwith the inlet of the heat exchanger; and a distribution system forreceiving reformed hydrocarbon fuel from the heat exchanger.

While this disclosure has been described as having exemplaryembodiments, the present disclosure can be further modified within thespirit and scope of this disclosure. This application is thereforeintended to cover any variations, uses, or adaptations of the disclosureusing its general principles. Further, this application is intended tocover such departures from the present disclosure as come within knownor customary practice in the art to which this disclosure pertains andwhich fall within the limits of the appended claims.

What is claimed is:
 1. A gas turbine engine, the gas turbine enginecomprising: a compressor, a combustor, and a turbine in serial flowrelationship; a heat exchanger having an inlet, an outlet, and aninternal surface coated with a catalyst, the heat exchanger beinglocated upstream of the compressor; a source of hydrocarbon fuel influid communication with the inlet of the heat exchanger; a source ofoxygen in fluid communication with the inlet of the heat exchanger; anda distribution system for receiving reformed hydrocarbon fuel from theheat exchanger.
 2. The gas turbine engine of claim 1, wherein thedistribution system delivers reformed hydrocarbon fuel to the gasturbine engine.
 3. The gas turbine engine of claim 1, wherein the gasturbine engine includes a hyperburner or augmentor.
 4. The gas turbineengine of claim 3, wherein the distribution system delivers reformedhydrocarbon fuel to the hyperburner or augmentor.
 5. The gas turbineengine of claim 3, wherein the hyperburner or augmentor includes ahyperburner fuel nozzle, and wherein the distribution system isconfigured to selectively deliver liquid hydrocarbon fuel or reformedhydrocarbon fuel to the fuel nozzle.
 6. The gas turbine engine of claim5, wherein the distribution system delivers liquid hydrocarbon fuel tothe combustor and reformed gaseous hydrocarbon fuel to the hyperburner.7. The gas turbine engine of claim 1, wherein the hydrocarbon fuelincludes an aircraft jet fuel, a gasoline, a kerosene, a rocketpropulsion fuel, a diesel fuels, or blends, mixtures, or combinationsthereof.
 8. The gas turbine engine of claim 1, wherein the source ofoxygen includes an oxygen-containing species.
 9. The gas turbine engineof claim 8, wherein the oxygen-containing species comprises steam orliquid water.
 10. The gas turbine engine of claim 1, wherein the sourceof oxygen includes an oxygenated fuel.
 11. The gas turbine engine ofclaim 8, wherein the oxygenated fuel comprises ethanol or methanol. 12.The gas turbine engine of claim 1, wherein the system includes aplurality of heat exchangers or heat exchanger branches.
 13. The gasturbine engine of claim 12, wherein the plurality of heat exchangers arearranged in series.
 14. The gas turbine engine of claim 12, wherein theplurality of heat exchangers are arranged in parallel.
 15. A method ofoperating a gas turbine engine, the gas turbine engine having acompressor, a combustor, and a turbine in serial flow relationship, thegas turbine engine also having a heat exchanger upstream of thecompressor, the heat exchanger being in fluid communication with asource of hydrocarbon fuel and a source of water and having an internalsurface coated with a catalyst, the method comprising: introducing ahydrocarbon fuel into the heat exchanger; introducing oxygen into theheat exchanger; contacting the hydrocarbon fuel with the catalyst; andcracking the hydrocarbon fuel to form a reformed hydrocarbon fuel andremove heat from the aircraft component.
 16. The method of claim 15,wherein the hydrocarbon fuel includes an aircraft jet fuel, a gasoline,a kerosene, a rocket propulsion fuel, a diesel fuels, or blends,mixtures, or combinations thereof.
 17. The method of claim 15, whereinthe source of oxygen is an oxygen-containing species or an oxygenatedfuel.
 18. The method of claim 15, wherein the method includes aplurality of heat exchangers.
 19. The method of claim 18, wherein theplurality of heat exchangers are arranged in series or in parallel. 20.The method of claim 18, wherein a distribution system delivers reformedfuel to a gas turbine engine.